This invention relates to a method for cooling an airplane engine operated with cryogenically stored propellant by means of cooling air.
The temperatures occurring in an airplane engine for high-speed airplanes (hypersonic airplanes) are extremely high because of the high Mach number. This requires the removal of an extremely high amount of heat in order to maintain the material temperature of the components in the hot are of the engine within permissible technical limits. It is an additional problem that, comparatively, the air temperatures at the air intake of a hypersonic engine, particularly one in hypersonic flight operation and thus exclusively switched-on ramjet propulsion, are extremely high (1,700.degree. C. and higher). Attempting to thermodynamically cool the even hotter engine components in the area of the combustion chamber as well as the turbine and parts connected behind them, such as afterburner elements, by correspondingly heated air cannot be implemented in practice. There is an additional problem in that cooling of the turbine, in particular a high-pressure turbine of the basic turbo-engine, can only take place by means of compressed air which must have approximately the same pressure as the pressure existing in the high-pressure turbine area.
Therefore, in the case of previous cooling concepts such as, for example, conventional gas turbine engines, the cooling air for the turbine is branched off from the compressor. This air, which is heated by the compression process, is fed to a cooling air cooler where it is intermediately cooled while utilizing the cooling capacity of the propellant which is carried along in a liquid state but is burned in a gaseous state. Hydrogen is particularly suitable as a propellant in this case.
This known solution has a number of disadvantages and problems. Since, for an effective cooling, approximately 10 to 20% of the total air mass flow must be branched off behind the compressor for cooling purposes, the cooling air cooler has large structural dimensions and hence a high weight. This is intensified in view of the minimizing of the air-side pressure losses in the heat exchanger. The reason for this is that the pressure loss along the total cooling air path from the compressor to the turbine must not be higher than that of the main flow in order to ensure a sufficient air flow and thus a satisfactory cooling. A blowing-out at the leading edge of the first turbine stator must also be ensured, where not only the highest static pressure but also the highest temperature occurs in the turbine. Finally, under certain flying conditions, there is the general risk of icing problems.
From "Proceedings of the 1st Int. Conference on Hypersonic Flight in the 21st Century", Grand Fox USA, 1988, Page 125 and on, it is known to provide an air breathing rocket engine with a condenser which, using the carried-along liquid hydrogen, liquifies the air required for burning this hydrogen.
On the basis of the above-described problems, there is needed a method and apparatus of the above-described type which avoids the above-mentioned problems and permit an improvement of the efficiency and of the cooling. The apparatus should also provide a constructive simplification of the cooling air cooling system.
According to the present invention, these needs are met by operating the airplane engine using a cryogenically stored propellant, taking in cooling air from an outside environment, liquifying said cooling air via a heat exchange with the cryogenically stored propellant to obtain liquified cooling air, increasing the pressure of the liquified cooling air, and supplying the liquified cooling air to the components of the airplane engine to be cooled.
The definition of "environment" in this case means that the cooling air was not compressed in the compressor but flows in from the atmospheric environment directly by way of an opening in the airplane fuselage or engine.
The method according to the present invention is particularly suitable for so-called hypersonic engines, i.e. those which accelerate airplanes to multiple sonic speed. At these speeds, the air at the engine intake already has high temperatures because of the air ram. However, this method is also suitable for conventional engines if cryogenically stored propellant is used.
The essential advantages of this cooling concept are that the pump delivery for the pressure increase of the air in the liquid state is considerably lower than the required compressor delivery for the gaseous air. This amounts to approximately 1/200 of the delivery required for compressing the gaseous air. In addition, the lowering of the cooling air temperatures results in a decrease of the cooling air requirement. Thus, the cooling air requirement can be lowered to approximately half of its previous values, i.e. 5-10% of the air mass flow. In addition, since this amount of air does not flow through the compressor, the structural dimensions of the compressor and therefore the compressor weight may be significantly decreased.
Because of the reduced cooling air flow, it is also advantageous that a smaller enthalpy gradient will occur in the turbine. Thus, in turn, leads to an increased pressure ratio in the nozzle and therefore to an increased thrust. The reason for this is that the reduction of the delivery requirement of the compressor outweighs the decrease of the turbine mass flow caused by the lower cooling air requirement (and thus the air mass flow).
Finally, it is an advantage of the present invention that the compressor intake surface, which determines the aerodynamic frontal drag of the airplane engine, can be decreased by the advantageous mass flow reduction. This has a considerable effect particularly in the case of high flight Mach numbers.
Hydrogen is preferably used as the propellant which is stored cryogenically, thus at very low temperatures, and is transported in the airplane. However, any other propellant which can be stored cryogenically, such as methane, is also suitable for this purpose. The propellant may be cooled to such an extent that it is present in the tank in a partially solidified state of aggregation in the form of slush.
When hydrogen (H.sub.2) is used, the temperature of the propellant which is carried along is approximately 20-40K., and the temperature during the injection into the combustion chambers is up to 1,000K. When hydrogen is used, the propellant, at the start of the heat exchange and thus while entering into the condenser, is preferably gaseous because the high specific heat capacity of the hydrogen is utilized. However, as an alternative, it is also possible to first evaporate the propellant in the condenser and thus utilize the evaporation enthalpy of the propellant.
In an advantageous further embodiment of the invention, the cooling air is branched off in front of the compressor of the airplane engine. In particular, air is branched off from the intake boundary layer of the engine. This advantageously reduces the installation losses in the area of the engine intake, and the intake cross-section may be reduced.
In an alternative further embodiment of the invention, the cooling air is branched off from the fuselage boundary layer of the airplane. As a result, the intake cross-section of the airplane engine can advantageously be reduced. Further, a duct, which up to now had been required for the removal of the fuselage boundary layer, is no longer required or may be reduced. As a result, this arrangement permits an advantageous use of the boundary layer flow which up to now had only caused losses and cooling problems in the area of the duct.
The apparatus of the present invention includes an air supply line for the supply of cooling air. The air supply line has a connection to the environment and is connected with a condenser. The liquified cooling air can be conveyed from the condenser by way of a pump and a cooling air line. The liquified cooling air can be in a liquid or vapor state and conveyed to the components to be cooled. In this case, the condenser can be driven by means of cryogenically stored propellant which can be fed by way of an inflow line and can be removed by way of an outflow line.
The resulting advantages correspond essentially to those described above concerning the method.
The condenser is preferably constructed as a countercurrent heat exchanger or as a cross-type/countercurrent heat exchanger in order to achieve a rate of exchange that is as high as possible.
Other objects, advantages and novel features of the present invention will become apparent from the following detailed description of the invention when considered in conjunction with the accompanying drawings.